FIG. 3 schematically shows a rear fuselage part A of such an aircraft fuselage F, made up of a number of fuselage parts, to explain the general problem.
The rear fuselage part A, shown in FIG. 3, of the aircraft fuselage F is delimited in the direction of the end of the fuselage by the pressure dome 1, which is curved towards the rear. The fuselage part A comprises six segments, which together form a cylinder.
Reference numeral 2 designates a floor, which is fitted on corresponding transverse beams 2a. The outer skin or casing 5 is fitted on annular frames 3. To stiffen the skin of the fuselage, stringers 4 are provided between the frames 3, and these stringers 4 connected to the outer skin 5, for example by adhesive bonding or riveting.
The stringers 4 are usually Z-, L- or I-shaped longitudinal stiffening elements, which provide a second load path in the event of a damaged skin (large damage capability). The stringers 4 run perpendicularly in relation to the frames and consequently parallel to the longitudinal axis of the aircraft. The stringers 4 are conventionally produced from an aluminium alloy.
Recently, a change has taken place in favour of using fibre-metal laminate structures (FML) for the outer skin 5 instead of the original technique of producing it from monolithic aluminium structures. Examples of such laminates are disclosed by WO 94/01277.
GLARE® is a laminate-like material combination that comprises a plurality of layers, each only a few tenths of a millimeter thick. These layers alternately consist of aluminium and a glass fibre laminate and are adhesively bonded under pressure. The word GLARE is an acronym for “Glass Fibre Reinforced Aluminium”. It was developed especially for aircraft construction and used for the first time over a large area in the AIRBUS A 380, in which large parts of the upper outer skin consist of GLARE. The advantages over aluminium are, in particular, its high damage tolerance, low density and good fire endurance. Fatigue cracks are bridged by the glass fibre layers, so that the crack propagation rate remains constantly low, irrespective of the length of crack, whereas in the case of aluminium the crack propagation rate increases sharply.
The density of GLARE is 9.5% to 13% below that of the aluminium customary in aircraft construction. In the case of GLARE, the glass fibres usually make up about 30% of the laminate. Since it is possible on account of the special properties of GLARE to reduce the skin thickness of the outer skin 5, i.e. to reduce the cross-sectional area of the outer skin 5, GLARE brings with it a considerable weight saving potential.
A disadvantage of GLARE is the reduced modulus of elasticity, which is around 57 GPa, in comparison with 70 GPa for aluminium. On account of the lower rigidity, there may be a shift in load from the GLARE components to other, neighbouring components. As a result, a weight advantage of the GLARE structure may be offset by a weight disadvantage of the surrounding structure. This is essentially the case for the fuselage structure over the centre-wing box, since the centre-wing box itself has a higher rigidity.
EP 1 336 469 A1 discloses a stringer for an aircraft or spacecraft that has an FML structure.